The Bell Jetranger 206B (III) helicopter was engaged in aerial
firefighting operations utilising an external water bucket and
staging out of a nearby national park camp ground. The pilot
reported that he started flying at approximately 0750 EST after
completing a pre-flight check of the helicopter which included
draining the fuel sump, inspecting the fuel and confirming 106 L or
27.9 United States Gallons (USG) of total indicated fuel. At
approximately 0825, while engaged in water bucket operations, he
discussed his fuel status with other company pilots on a common
radio frequency and noted 38 L (10 USG) of indicated fuel
remaining. He finished a swath run of the fire area, dropping water
and then decided to complete one more swath run before returning to
refuel.
Approaching the fire line, the helicopter entered a left turn at
approximately 200 ft above ground level (AGL). The pilot reported
that the helicopter was buffeted by strong turbulence, which caused
the helicopter to yaw left and go out of trim. He reported that the
engine power then began surging and, subsequently, an engine
flameout occurred. He continued the left turn, jettisoned the water
and initiated a power-off autorotation to a heavily wooded
area.
During the autorotation, the helicopter's main rotor blades
contacted nearby trees, damaging the blades and displacing the main
rotor transmission. The helicopter then came to rest on its right
side, on a 45 degree slope. The pilot exited the helicopter and was
picked up by another company helicopter and transported to hospital
for observation. The pilot suffered minor injuries and the
helicopter was substantially damaged. There was no evidence of any
in-flight or post-accident fire.
The pilot reported that he had no recollection of any
illuminated caution advisory lights during or prior to the event,
or of the position of the switches and circuit breakers of the fuel
boost pumps. He stated that he followed flight manual checklists
when starting the helicopter and preparing for takeoff.
Wreckage examination
The helicopter sustained damage to the main rotor blades and
controls, main rotor transmission, transmission to engine
driveshaft, tail rotor driveshaft, pilot's perspex and right side
forward fuselage. Examination of the engine revealed no external
damage or damage to the compressor.
Fuel system examination and background
After the helicopter was repositioned upright, the fuel gauge
indicated approximately 19 L (5 USG). The fuel in the fuel tank was
examined and appeared to slightly cover the base of the fuel boost
pumps. Examination of the airframe fuel filter revealed full fuel
in the filter bowl and no contamination. Examination of the engine
fuel filter revealed a small amount of fuel and no contamination.
The forward fuel boost pump circuit breaker was noted as
disengaged.
Unusable fuel was defined as, `Fuel that cannot be used in
flight with wings level and at cruise angle of attack (or nose 3
degrees up)'. Of the total fuel on board, 4 L (1 USG) were
unusable. Upon removal from the accident site, the helicopter's
fuel system was drained and 23.5 L (6.2 USG) of fuel were removed.
The fuel correction card annotated, `gauge indicates 38 L (10 USG)
for an actual of 45.6 L (12 USG)'. The fuel correction card values
were verified by adding measured amounts of fuel.
An option on that model helicopter was a FUEL LOW caution
advisory light that illuminated with approximately 76 L (20 USG) of
total fuel remaining. The helicopter was not equipped with that
advisory light. It was equipped with two electrically operated
submerged fuel boost pumps located in the fuel cell and connected
in parallel to the engine's fuel supply line. Those pumps were
located on the helicopter's centre-line. Both boost pumps were
examined following the accident and they were determined to be
serviceable.
Fuel consumption
According to the operator's operations manual, the fuel
consumption rate of the Bell 206B (III) helicopter was 110 L (28.9
USG) per hour. In the 35 minutes of operating time that the pilot
reported prior to the accident, approximately 64 L (16.8 USG) of
fuel should have been consumed, leaving approximately 30.8 L (8.1
USG) of useable fuel remaining. That amount should have been
sufficient for approximately 17 minutes of engine operating
time.
Engine auto-reignition
The helicopter was not fitted with an optional engine
auto-reignition system. Because of the low height AGL at the time
of the engine surging and flameout, the pilot did not have an
opportunity to attempt a manual restart of the engine.
Engine testing
The Rolls-Royce Allison model 250-C20B engine, serial number CAE
840551, was removed from the helicopter and transported to an
engine test cell for testing. Following motoring and priming, the
engine started on the first attempt with Turbine Operating
Temperatures, N1 (gas generator speed) and N2 (power turbine speed)
values within normal operating parameters. Engine deceleration and
acceleration tests were conducted in order to simulate power
changes experienced during flight. The results of the testing
indicated that the engine was serviceable at the time of the
accident.
Helicopter manoeuvring in turns and flight manual
requirements
During coordinated turns, centrifugal force acting on the
helicopter and fuel tank causes the fuel to collect evenly in the
bottom of the fuel tank. It is then available for pick up by the
fuel boost pumps at the boost pump inlets and through to the inlet
of the main fuel supply line and to the engine. During
uncoordinated turns, centrifugal force may not displace the fuel to
the bottom of the tank evenly and instead fuel sloshing may take
place.
The helicopter's flight manual contained a warning regarding
flight with one fuel boost pump inoperative. It stated, `Due to
possible fuel sloshing in unusual attitudes or out of trim
conditions and one or both fuel boost pumps inoperative, the
unusable fuel is ten [US] gallons'.
Component testing
The forward fuel boost pump circuit breaker was removed and
tested. Aircraft system 28 DC voltage was applied and varying
amperes were introduced to the circuit breaker in an attempt to
determine its serviceability. The circuit breaker operated normally
up to its rated 10 ampere rating with no anomalies noted.
Pilot's shoulder harness
The pilot's left shoulder harness had broken and separated at a
point just forward of and below the pilot's shoulder. The
manufacturer's date stamped on the harness belt was March 1973. The
pilot's seat belt had an inspection tag attached with the
inspection date 11/99 annotated. Details of the inspection were not
annotated in the helicopter's documentation.
Pilot shoulder harness testing
The pilot's left shoulder harness was sent to an independent
belt and harness testing and repair organisation for testing. The
webbing was identified as MIL-T-50368 Type IV, 2 inch Nylon
Webbing, rated at 2,000 pounds strength. The rated assembly
strength of the harness assembly was 1,500 pounds. Testing revealed
that the webbing failed at a value of 391 pounds, or less than 20
percent of the original strength of the material. Factors
contributing to the loss of original strength were ageing related
to ultraviolet light exposure, abrasion damage and contamination by
turbine oil.
Seat belt and shoulder harness standards
Australian Civil Aviation Order (CAO) Part 108, Section 108.42,
contained specifications for aircraft safety belts (seat belts),
harnesses (shoulder harnesses) and inertia reels manufactured in
Australia. Contained within CAO Part 108, Section 108.42 was a
reference to Technical Standard Order (TSO)-C22g, which specified
minimum performance standards of aircraft seat belts when new.
CAO Part 108, Section 108.42, also contained a reference to
British Air Registration Board (ARB)/Civil Aviation Authority (CAA)
Specification Number four (issue two). That document specified
minimum performance standards for British manufactured aircraft
shoulder harnesses when new and was not applicable to equipment
manufactured in the United States of America (USA) such as the
separated harness in this occurrence.
The Australian Civil Aviation Safety Authority (CASA)
Airworthiness Directive, AD/RES/29 amendment 1, mandated
identification of aircraft seat belts to indicate that the part was
an item approved by the manufacturer. Part of that directive
mandated appropriate identification to demonstrate compliance with
CAO Part 108, Section 108.42.
The helicopter was manufactured in 1979 in the US under the
requirements of the US Federal Aviation Administration (FAA),
Federal Aviation Regulations. FAA Technical Standard Order
(TSO)-C22g (amended 1993) outlined minimum performance standards
for aircraft seat belts when new. TSO-C114 (initial issue 1987)
outlined minimum performance standards for aircraft torso
restraints (shoulder harnesses) when new. Prior to 1987, there were
no requirements for shoulder harnesses. Only TSO-C22 was in effect
when the occurrence helicopter was manufactured.
At the time of the occurrence, there were no requirements to
confirm compliance to applicable design standards of shoulder or
seat belt harnesses while in service, or to specifically identify
shoulder harnesses.
